US9416661B2 - Axial compressor and associated operating method - Google Patents

Axial compressor and associated operating method Download PDF

Info

Publication number
US9416661B2
US9416661B2 US13/287,308 US201113287308A US9416661B2 US 9416661 B2 US9416661 B2 US 9416661B2 US 201113287308 A US201113287308 A US 201113287308A US 9416661 B2 US9416661 B2 US 9416661B2
Authority
US
United States
Prior art keywords
compressor
cooling medium
hollow space
stator blade
flow
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active, expires
Application number
US13/287,308
Other languages
English (en)
Other versions
US20120114459A1 (en
Inventor
Francois Benkler
Sascha Dungs
Harald Hoell
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens Energy Global GmbH and Co KG
Original Assignee
Siemens AG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens AG filed Critical Siemens AG
Assigned to SIEMENS AKTIENGESELLSCHAFT reassignment SIEMENS AKTIENGESELLSCHAFT ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: BENKLER, FRANCOIS, DUNGS, SASCHA, HOELL, HARALD
Publication of US20120114459A1 publication Critical patent/US20120114459A1/en
Application granted granted Critical
Publication of US9416661B2 publication Critical patent/US9416661B2/en
Assigned to Siemens Energy Global GmbH & Co. KG reassignment Siemens Energy Global GmbH & Co. KG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SIEMENS AKTIENGESELLSCHAFT
Active legal-status Critical Current
Adjusted expiration legal-status Critical

Links

Images

Classifications

    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02Blade-carrying members, e.g. rotors
    • F01D5/08Heating, heat-insulating or cooling means
    • F01D5/081Cooling fluid being directed on the side of the rotor disc or at the roots of the blades
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02Blade-carrying members, e.g. rotors
    • F01D5/08Heating, heat-insulating or cooling means
    • F01D5/085Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor
    • F01D5/088Heating, heat-insulating or cooling means cooling fluid circulating inside the rotor in a closed cavity
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/18Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/181Blades having a closed internal cavity containing a cooling medium, e.g. sodium
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/18Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/185Liquid cooling
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/18Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/186Film cooling
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12Blades
    • F01D5/14Form or construction
    • F01D5/18Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187Convection cooling
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00Stators
    • F01D9/02Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01DNON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00Stators
    • F01D9/06Fluid supply conduits to nozzles or the like
    • F01D9/065Fluid supply or removal conduits traversing the working fluid flow, e.g. for lubrication-, cooling-, or sealing fluids
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12Cooling of plants
    • F02C7/14Cooling of plants of fluids in the plant, e.g. lubricant or fuel
    • F02C7/141Cooling of plants of fluids in the plant, e.g. lubricant or fuel of working fluid
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12Cooling of plants
    • F02C7/16Cooling of plants characterised by cooling medium
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12Cooling of plants
    • F02C7/16Cooling of plants characterised by cooling medium
    • F02C7/18Cooling of plants characterised by cooling medium the medium being gaseous, e.g. air
    • F02C7/185Cooling means for reducing the temperature of the cooling air or gas
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00Details, component parts, or accessories
    • F04D29/05Shafts or bearings, or assemblies thereof, specially adapted for elastic fluid pumps
    • F04D29/053Shafts
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00Details, component parts, or accessories
    • F04D29/40Casings; Connections of working fluid
    • F04D29/52Casings; Connections of working fluid for axial pumps
    • F04D29/54Fluid-guiding means, e.g. diffusers
    • F04D29/541Specially adapted for elastic fluid pumps
    • F04D29/542Bladed diffusers
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04DNON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00Details, component parts, or accessories
    • F04D29/58Cooling; Heating; Diminishing heat transfer
    • F04D29/582Cooling; Heating; Diminishing heat transfer specially adapted for elastic fluid pumps
    • F04D29/584Cooling; Heating; Diminishing heat transfer specially adapted for elastic fluid pumps cooling or heating the machine
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00Components
    • F05D2240/20Rotors
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00Function
    • F05D2260/20Heat transfer, e.g. cooling
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00Function
    • F05D2260/20Heat transfer, e.g. cooling
    • F05D2260/201Heat transfer, e.g. cooling by impingement of a fluid
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00Function
    • F05D2260/20Heat transfer, e.g. cooling
    • F05D2260/202Heat transfer, e.g. cooling by film cooling
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05DINDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00Function
    • F05D2260/20Heat transfer, e.g. cooling
    • F05D2260/208Heat transfer, e.g. cooling using heat pipes

Definitions

  • the invention refers to an axial compressor, especially for a gas turbine, for compressing a flow medium, having a multiplicity of compressor stator blades, which are assembled to form stator blade rows and fastened in each case on a stator blade carrier, and having a multiplicity of compressor rotor blades, which are assembled to form rotor blade rows and fastened in each case on a compressor disk of a compressor shaft, wherein two consecutive compressor disks in each case enclose a hollow space lying between them, and wherein the last compressor disk, as seen in the flow direction of the flow medium, delimits a rear hollow space.
  • the invention furthermore refers to a method for operating such an axial compressor.
  • Gas turbines are used in many fields for driving generators or driven machines.
  • the energy content of a fuel is utilized for producing a rotational movement of a turbine shaft.
  • the fuel is combusted in a combustion chamber, wherein compressed air is supplied from an air compressor.
  • the air compressor in this case is customarily constructed as an axial compressor.
  • the operating medium, under high pressure and at high temperature, which is produced in the combustion chamber as a result of the combustion of the fuel is directed in this case through a turbine unit—which is connected downstream to the combustion chamber—where it expands, performing work.
  • the air compressor, or compressor for short, and the turbine unit are customarily arranged on a common shaft so that the turbine unit drives the compressor during operation.
  • the combustion chamber of the gas turbine can be constructed as a so-called annular combustion chamber, in which a large number of burners, which are circumferentially arranged around the turbine shaft, open into a common combustion space which is enclosed by a high temperature-resistant surrounding wall.
  • the combustion chamber is designed as an annular structure in its entirety.
  • a multiplicity of combustion chambers can also be provided.
  • turbine rotor blades For producing the rotational movement of the turbine shaft, in this case a number of turbine rotor blades, which are customarily assembled into blade groups or blade rows, are arranged on this shaft.
  • a turbine disk on which the turbine rotor blades are fastened by means of their blade root, is customarily provided for each turbine stage.
  • turbine stator blades For flow guiding of the operating medium in the turbine unit, moreover, turbine stator blades, which are connected to the turbine casing and assembled to form stator blade rows, are customarily arranged between adjacent rotor blade rows.
  • the air compressor of such a gas turbine in respect to construction, is customarily constructed similarly to the turbine unit and, in a configuration as an axial compressor, comprises a multiplicity of compressor stator blades, which are assembled to form stator blade rows and fastened in each case on a stator blade carrier, and a multiplicity of compressor rotor blades, which are assembled to form rotor blade rows and fastened in each case on a compressor shaft.
  • a plurality of compressor stages are provided.
  • stator The entirety of all the rotating parts of the gas turbine—especially the shaft and the rotor blades—are also referred to as a rotor, and the stationary parts—especially the casing and the stator blades—are also referred to collectively as a stator.
  • the compressor shaft is customarily assembled from a multiplicity of compressor disks which are arranged one behind the other, as seen in the axial direction, and are held together by means of a tie bolt, for example.
  • the compressor shaft continues, via a shaft intermediate piece, as the turbine shaft.
  • each of the compressor disks customarily carries the compressor rotor blades of a rotor blade row, which rotor blades, by means of blade roots, for example, are fastened in corresponding fastening grooves of the compressor disk.
  • a compressor disk can also carry a plurality of rotor blade rows.
  • Annular hollow spaces also referred to as a cavity in each case in the following text, are customarily provided between two consecutive compressor disks in each case, as a result of which the total weight of the compressor shaft is reduced in comparison to a completely solid type of construction.
  • the last compressor disk as seen in the flow direction of the flow medium, in a conventional type of construction has a top face or end face which points towards the subsequent turbine unit and together with other components, for example, delimits or at least partially encloses a rear hollow space—also referred to as a rear cavity—which is separated from the flow passage for the flow medium.
  • a rear hollow space also referred to as a rear cavity
  • a plurality of groups of hollow spaces can also be provided in the compressor shaft, wherein, for example, the hollow spaces of a first group lie further on the outside, as seen in the radial direction, whereas the hollow spaces of a second, and, if applicable, of further groups, lie further on the inside.
  • the air in the compressor should be compressed as intensely as possible. Contingent upon the pressure in the compressor which increases more and more along the flow direction of the gas, the temperature at the compressor exit also increases along with it. The maximum permissible operating temperature of the material of the rear compressor disks is possibly reached in the process.
  • the maximum permissible operating temperature for available materials represents a limiting boundary condition for the development of gas turbines with regard to the compressor exit temperature. If there is the risk of this limit being exceeded, for example in the case of high ambient temperature, the operating mode of the machine must be throttled. As a result, the potential of the gas turbine cannot be fully utilized.
  • the last compressor disk as seen in the flow direction of the flow medium, especially in its top region or end region, is cooled by means of impingement with cooling air.
  • a cooling-air cooler for example, which serves essentially for the supply of the front turbine blading with cooled-down cooling air.
  • This cooling air is fed into the rotor through the so-called shaft cover, specifically a shaft cover or casing arranged downstream of the air compressor, as seen in the flow direction of the flow medium. From there, the cooling air then finds its way into the turbine blading. Some of this cooling air is directed from the shaft cover into the cavity downstream of the last compressor disk for cooling the top region of the last compressor disk in the process.
  • the cooling air for the cavity travels a relatively long way through various components of the gas turbine, specifically first through the feed line to the shaft cover and then through the shaft cover itself, which are exposed to circumflow by hot compressor air.
  • the temperature of the cooling air significantly increases before it reaches the cavity, as a result of which the cooling potential for the last compressor disk is greatly reduced.
  • the last compressor disk is cooled only from one side, and the compressor disks which lie further upstream of it, as seen in the flow direction of the flow medium,—which admittedly are not quite as high as the last compressor disk but are certainly appreciably thermally loaded—are possibly not cooled at all.
  • the invention is therefore based on the object of further developing an axial compressor of the type referred to in the introduction in such a way that with means which are kept simple an effective cooling of the rear compressor disks, especially of the last compressor disk, as seen in the flow direction of the flow medium, is achieved. Furthermore, a corresponding operating method is to be disclosed.
  • this object is achieved according to the invention by at least one cooling medium feed duct which leads through a stator blade carrier and through a compressor stator blade which is fastened on the stator blade carrier and arranged upstream of the last compressor disk, as seen in the flow direction of the flow medium, which cooling medium feed duct, at the tip end of the compressor stator blade facing the compressor shaft, via a discharge opening arranged there, opens into one of the hollow spaces lying between the compressor disks, wherein this hollow space, via at least one cooling medium transfer passage which is led through the subsequent compressor disks, is connected to the rear hollow space.
  • the invention in this case is based on the consideration that particularly efficient cooling of the rear compressor disks, especially of the last compressor disk, as seen in the flow direction of the flow medium, can be achieved by it being impinged upon by cooling medium of comparatively low temperature on at least one of its two top faces or end faces, preferably on both top faces or end faces. In this case, excessive heating of the cooling medium, which is introduced into the adjacent hollow space, on its way into the hollow space should be avoided.
  • the cooling medium is fed to a hollow space which adjoins or is adjacent to the compressor disks which are to be cooled via at least one of the compressor stator blades of the stator blade row which lies directly opposite the hollow space in the radial direction. Since this stator blade row and the hollow space which is to be cooled are located essentially at the same position, as seen in the axial direction, the feed of cooling medium is carried out essentially in the radial direction from the outside inwards so that comparatively long transporting or feed paths in the axial direction through various hot components of the gas turbine, in which an undesirable heating of the cooling medium could take place, are avoided.
  • the introduction of the cooling medium is therefore first carried out via the compressor stator blades into one of the front hollow spaces which lie between two compressor disks, for example into the hollow space which lies upstream of the last compressor disk, as seen in the flow direction of the flow medium, and therefore usually upstream of the last rotor blade row.
  • the cooling medium via at least one cooling medium transfer passage, which is subsequently also referred to as transfer passage for short and is led through the subsequent compressor disk(s), is then directed into the rear hollow space downstream of the last compressor disk so that this region is also cooled with the components adjacent to it.
  • the last compressor disk is therefore cooled on both sides, i.e. on both top faces or end faces.
  • the “used” cooling medium can finally discharge as a result of gap leakage, or in some other way, into the flow passage for the flow medium which is to be compressed and/or into downstream cooling medium discharge passages or the like.
  • the respective transfer passage can be advantageously introduced into the last compressor disk in the style of an axial bore for this purpose.
  • the cooling medium is introduced via the compressor stator blades into a hollow space which is located further forward in the compressor shaft and is then directed into the rear hollow space via a number of hollow spaces which are connected in between in respect to flow, corresponding transfer passages through all the compressor disks lying in between are to be expediently provided.
  • Cooling air which for example is extracted as partial flow from the compressor air flow further upstream in a colder region of the compressor, is expediently used as cooling medium.
  • recooling of the cooling air by means of external cooling-air coolers or the like can also be provided.
  • a plurality of transfer passages which are especially connected in parallel on the cooling medium side, are expediently provided for each of the compressor disks in question, as seen in the circumferential direction.
  • the respective transfer passage in this case can be led through the corresponding compressor disk especially in the style of an axial bore.
  • stator blade carrier or in an encompassing casing component for an annular cooling medium distribution chamber, to which are connected the sections of the cooling medium ducts which are arranged in the compressor stator blades.
  • the compressor stator blades of the stator blade row which are provided for feed of cooling medium are connected at their tip end to a common annular body which on one side, by its outside face, delimits a flow passage for the flow medium which is comparatively hot at this point, and which on the other side, by its inside face, delimits the hollow space which is to be cooled and into which the introduction of cooling medium is carried out.
  • the annular body which is constructed in the style of a shroud, therefore seals the hollow space in relation to the flow passage and thermally isolates both space regions from each other.
  • the respective cooling medium feed duct is expediently led through the annular body so that the cooling medium discharge opening is located on its inside face which faces the hollow space.
  • the rear hollow space on the side lying opposite the rear end face of the last compressor disk, can be delimited by an end face of, for example, an annular or hollow cylindrical rotor cover (shaft cover).
  • the end face of the rotor cover in this case is also cooled by means of the cooling medium which transfers into the rear hollow space via the transfer passage during operation.
  • the use of the described axial compressor as an air compressor in a gas turbine is especially preferred, wherein the air compressor and the turbine unit of the gas turbine are advantageously arranged along a common shaft.
  • the axial compressor it is also naturally conceivable to operate the axial compressor as a stand-alone unit for other application purposes in which a flow medium is to be compressed. Said advantages of the improved cooling of the compressor exit region come into effect in this case also.
  • the object which is refereed to in the introduction is achieved by a cooling medium being introduced into one of the hollow spaces lying between the compressor disks by means of at least one cooling medium feed duct which is led through a stator blade carrier and through a compressor stator blade which is fastened on the stator blade carrier and arranged upstream of the last compressor disk, as seen in the flow direction of the flow medium, wherein from there the cooling medium is introduced into the rear hollow space via at least one cooling medium transfer passage which is led through the subsequent compressor disks.
  • the advantages which are associated with the invention are especially that in an axial compressor, by means of a feed of cooling medium which is carried out essentially in the radial direction via a spatially adjacent stator blade row, effective cooling of the thermally especially loaded rear compressor disks and of the adjacent components is made possible.
  • a lower thermal loading of the rear compressor disks or a possible increase of the compressor exit pressure with an unchanged loading of the compressor disks results from this.
  • FIG. 1 shows a half-section through a gas turbine
  • FIG. 2 shows an enlarged detail of the compressor of the gas turbine according to FIG. 1 with a previously provided cooling device for cooling the last compressor disk, and
  • FIG. 3 shows a detail from FIG. 1 with an improved cooling device, compared with FIG. 2 , for cooling the rear compressor disks.
  • the gas turbine 1 has a compressor 2 for combustion air, a combustion chamber 4 and also a turbine unit 6 for driving the compressor 2 and for driving a generator or a driven machine, which is not shown.
  • the turbine unit 6 and the compressor 2 are arranged on a common turbine shaft 8 , which is also referred to as a turbine rotor, to which the generator or the driven machine is also connected, and which is rotatably mounted around its center axis 9 .
  • the combustion chamber 4 which is constructed in the style of an annular combustion chamber, is equipped with a number of burners 10 for combusting a liquid fuel or gaseous fuel.
  • the turbine unit 6 has a number of rotatable turbine rotor blades 12 which are connected to the turbine shaft 8 .
  • the turbine rotor blades 12 are arranged on the turbine shaft 8 in a ring-like manner and therefore form a number of rotor blade rows.
  • the turbine unit 6 comprises a number of stationary turbine stator blades 14 which are also fastened in a ring-like manner on a stator blade carrier 16 of the turbine unit 6 , forming stator blade rows.
  • the turbine rotor blades 12 in this case serve for driving the turbine shaft 8 as a result of impulse transfer from the operating medium M which flows through the turbine unit 6 .
  • the turbine stator blades 14 serve for flow guiding of the operating medium M between two consecutive rotor blade rows or rotor blade rings in each case, as seen in the flow direction of the operating medium M.
  • a consecutive pair consisting of a ring of turbine stator blades 14 , or a stator blade row, and of a ring of turbine rotor blades 12 , or a rotor blade row, in this case is also referred to as a turbine stage.
  • the compressor 2 of the gas turbine 1 is constructed similarly to the turbine unit 6 . It comprises a multiplicity of compressor rotor blades 18 , which are assembled to four' rotor blade rows and by their blade roots are fastened on the turbine shaft 8 , referred to as the compressor shaft 20 in this section of the gas turbine 1 , which compressor rotor blades project into a flow passage 22 for the inducted flow medium S, in this case air.
  • the compressor rotor blades 18 which are set in rotation via the compressor shaft 20 , perform compression work on the flow medium S and transport it in the direction towards the turbine unit 8 .
  • the stationary compressor stator blades 24 which are assembled to faun stator blade rows, on the other hand, serve for flow guiding of the flow medium S between two consecutive rotor blade rows in each case, as seen in the flow direction of the flow medium.
  • the compressor stator blades 24 are fastened on associated stator blade carriers 26 which in their turn are connected, in a way not shown, to the outer compressor casing and which—possibly together with further ring segments—form the outer boundary of the flow passage 22 .
  • the stator blade carriers 26 can be assembled from a plurality of segments.
  • FIG. 2 shows in an enlarged view the exit region or discharge region of the compressor 2 and the subsequent transition region, in the flow direction 28 of the flow medium S, to the combustion chamber 4 and to the turbine unit 6 .
  • the compressor shaft 20 is assembled from a plurality of compressor disks 30 , which are arranged one behind the other in a stacked manner, of which only a single compressor disk 30 , specifically the rear or last compressor disk 30 , as seen in the flow direction 28 of the flow medium S, is visible in FIG. 2 .
  • the respective compressor disk 30 carries the compressor rotor blades 18 of the associated rotor blade row on its periphery.
  • the circumferential surface of the respective compressor disk 30 in the region of its extent, at the same time forms the inner boundary of the flow passage 22 .
  • the inner boundary of the flow passage 22 is formed in each case by the outer side 32 of an annular body 34 which is connected to the tip ends 36 of the compressor stator blades 24 of the associated stator blade row.
  • the annular gap 38 which is located between the respective—spatially fixed—annular body 34 and the axially adjacent—rotating—compressor disk 30 , can be sealed in a conventional manner, by means of a labyrinth seal 40 , for example.
  • the last compressor stage downstream of the last rotor blade row comprises two directly consecutive stator blade rows (so-called double row arrangement), with which is associated a common annular body 34 . It does not depend upon this detail, however, in the present case.
  • the flow passage 22 widens for the flow medium S, which is compressed in the compressor 2 , in the style of a diffuser.
  • the inner boundary of the flow passage 22 is formed in this region by means of the circumferential surface of an annular, so-called shaft cover 42 .
  • the stationary shaft cover 42 encloses the rotating turbine shaft 8 which, as an extension of the compressor shaft 20 , extends towards the turbine unit 6 , and can be assembled from individual shaft segments 44 or disks.
  • the shaft cover 42 extends in the axial direction towards the compressor 2 almost as far as the annular body 34 of the last (double) stator blade row.
  • the shaft cover 42 On the end face which is oriented towards the compressor 2 , the shaft cover 42 has an annular flange 46 with an end face 48 which is at a distance from the annular body 34 by means of an axial annular gap 50 .
  • the annular flange 46 is at a distance towards the turbine shaft 8 by means of a further, in this case radial, annular gap 52 .
  • the annular gap 50 can be provided with suitable sealing means in order to prevent a transfer of the comparatively hot flow medium S from the flow passage 22 at the compressor outlet into the hollow space 54 , also referred to as a cavity, which is delimited or enclosed by the last compressor disk 30 , the annular body 34 , the annular flange 46 and the corresponding section of the turbine shaft 8 .
  • the components which are adjacent to the hollow space 54 can be exposed to a considerable thermal load during operation of the gas turbine 1 or of the compressor 2 .
  • a cooling medium K in this case cooling air
  • the cooling medium K in this case is directed, via a cooling medium feed line 56 , into a cooling medium duct 58 of, for example, cylindrical contour, which is integrated into the shaft cover 42 .
  • the cooling medium flows through one or more transfer passages 60 , which are introduced into the annular flange 46 , into the hollow space 54 so that the desired cooling of the last compressor disk 30 is realized in the manner of an impingement cooling on the end face 62 .
  • the discharge of the “used” cooling medium K is carried out by means of gap leakage, for example, at the annular gaps 38 , 50 and 52 .
  • each of the two rear compressor disks 64 and 66 carries two stator blade rows, which in this case, however, is not of vital importance.
  • a stator blade row is located between two rotor blade rows in each case and downstream of the last rotor blade row, as seen in the flow direction of the flow medium S.
  • a stator blade row specifically the cooling medium feed-stator blade row 68 which is to be described in more detail further down, is located in the axial region between the last but one compressor disk 64 and the last compressor disk 66 .
  • annular hollow space 70 which is delimited by the rear end face 72 of the last but one compressor disk 64 and by the front end face 74 of the last compressor disk 66 .
  • the rear end face 76 of the last compressor disk 66 delimits, at least partially, an annular rear hollow space 78 .
  • the hollow spaces 70 and 78 are located approximately at the same position, as seen in the radial direction.
  • a feed of cooling medium K is carried out via a number of compressor stator blades 24 of the cooling medium feed-stator blade row 68 .
  • the cooling medium feed duct 80 in an also radially extending end section 88 , continues inside the annular body 90 , which is connected to the compressor blade 24 , and terminates in a discharge opening 94 on the inner side 92 of the annular body 90 facing the hollow space 70 .
  • the annular body 90 seals the hollow space 70 in relation to the flow passage 22 , wherein the outer side 96 of the annular body 90 in the region of its extent forms the inner boundary of the flow passage 22 .
  • Suitable shaft seals for example in the form of a labyrinth seal, can be arranged in the annular gap 98 which is formed between the annular body 90 and the adjacent compressor disks 64 and 66 .
  • the cooling medium K which is fed to the hollow space 70 is only slightly heated and can therefore develop a comparatively high cooling potential, especially when cooling the rear end face 72 of the last but one compressor disk 64 and the front end face 74 of the last compressor disk 66 .
  • the slightly heated cooling medium K via one or more cooling medium transfer passages 100 which are preferably arranged in a circumferentially distributed manner, is then directed into the rear hollow space 78 downstream of the last compressor disk 66 .
  • the respective cooling medium transfer passage 100 is introduced into the last compressor disk 66 in the style of an axial bore.
  • the respective axial bore connects the inlet opening 102 of the transfer passage 100 , which is arranged in the front end face 74 of the last compressor disk 66 , to the discharge opening 104 which is arranged in the rear end face 76 .
  • a feed of cooling medium, which is approximately uniform in the circumferential direction, into the hollow space 70 , which is designed as an annulus, is carried out via a plurality of the compressor stator blades 24 of the corresponding stator blade row 68 , wherein, for example, every compressor stator blade 24 , every other compressor stator blade or every third compressor stator blade, etc., of this stator blade row 68 , as seen in the circumferential direction, can be provided with a corresponding section of a cooling medium feed duct 80 and with a corresponding discharge opening 94 for cooling medium K.
  • duct sections are therefore connected in parallel on the cooling medium side and are fed simultaneously with fresh cooling medium K via a circumferentially extending cooling medium distribution chamber 106 , for example, which is shown only schematically in FIG. 3 and arranged in the stator blade carrier 26 or in an adjacent casing component.
  • FIG. 3 It is understood that some of the details shown in FIG. 3 have characters which are only by way of example. Deviating from the view which is selected here, for example the orientation of the end section 88 of the cooling medium feed duct 80 and the position of the discharge opening 94 can vary. Also, the feed of cooling medium could alternatively or additionally be carried out, via an associated stator blade row lying upstream of the stator blade row 68 , into a hollow space which lies upstream of the hollow space 70 , for example into the hollow space 108 , as seen in the flow direction of the flow medium S. In the case of this variant, the front hollow space 108 , which in such a way is exposed to admission of cooling medium K, in its turn, is connected via additional cooling medium transfer passages (not shown in FIG. 3 ) indirectly—i.e. via hollow spaces lying further to the rear—or directly to the rear hollow space 78 .

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Physics & Mathematics (AREA)
  • Thermal Sciences (AREA)
  • Fluid Mechanics (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US13/287,308 2010-11-04 2011-11-02 Axial compressor and associated operating method Active 2033-09-20 US9416661B2 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
EP10189974 2010-11-04
EP20100189974 EP2450531B1 (de) 2010-11-04 2010-11-04 Axialverdichterkühlung
EPEP10189974 2010-11-04

Publications (2)

Publication Number Publication Date
US20120114459A1 US20120114459A1 (en) 2012-05-10
US9416661B2 true US9416661B2 (en) 2016-08-16

Family

ID=43929085

Family Applications (1)

Application Number Title Priority Date Filing Date
US13/287,308 Active 2033-09-20 US9416661B2 (en) 2010-11-04 2011-11-02 Axial compressor and associated operating method

Country Status (4)

Country Link
US (1) US9416661B2 (ja)
EP (1) EP2450531B1 (ja)
JP (1) JP5411233B2 (ja)
CN (1) CN102454480B (ja)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11746675B2 (en) 2021-11-23 2023-09-05 Rolls-Royce Corporation Vane ring assembly for a gas turbine engine with dedicated through-flow vanes

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP5868802B2 (ja) * 2012-07-20 2016-02-24 株式会社東芝 タービン
US10480533B2 (en) * 2013-09-10 2019-11-19 United Technologies Corporation Fluid injector for cooling a gas turbine engine component
EP2868865A1 (de) * 2013-10-31 2015-05-06 Siemens Aktiengesellschaft Gasturbine sowie Verfahren zu deren Kühlung
DE102015219022A1 (de) * 2015-10-01 2017-04-06 Rolls-Royce Deutschland Ltd & Co Kg Strömungsleitvorrichtung und Turbomaschine mit mindestens einer Strömungsleitvorrichtung
US20170130732A1 (en) * 2015-11-06 2017-05-11 United Technologies Corporation Compressor exit seal
BE1026455B1 (fr) * 2018-07-09 2020-02-03 Safran Aero Boosters Sa Compresseur de turbomachine
JP2021032224A (ja) * 2019-08-29 2021-03-01 三菱パワー株式会社 圧縮機、ガスタービン
RU2735972C1 (ru) * 2020-05-10 2020-11-11 Владимир Дмитриевич Куликов Система воздушно-жидкостного охлаждения лопаток ступеней турбины турбореактивного двигателя
US11566532B2 (en) * 2020-12-04 2023-01-31 Ge Avio S.R.L. Turbine clearance control system

Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE3606597C1 (de) 1986-02-28 1987-02-19 Mtu Muenchen Gmbh Schaufel- und Dichtspaltoptimierungseinrichtung fuer Verdichter von Gasturbinentriebwerken
EP0735255A1 (en) 1995-03-31 1996-10-02 General Electric Company Compressor rotor cooling system for a gas turbine
JPH11125199A (ja) 1997-10-22 1999-05-11 Mitsubishi Heavy Ind Ltd 空気圧縮機ディスクの冷却によるクリアランス制御方法
JPH11315800A (ja) 1998-04-30 1999-11-16 Toshiba Corp 空気圧縮機
US6099244A (en) * 1997-03-11 2000-08-08 Mitsubishi Heavy Industries, Ltd. Cooled stationary blade for a gas turbine
EP1074694A2 (en) 1999-08-04 2001-02-07 General Electric Company Apparatus and methods for cooling rotary components in a turbine
CN1295184A (zh) 1999-11-05 2001-05-16 株式会社日立制作所 燃气涡轮、燃气涡轮机及用于其动叶片的冷却剂收集法
US6250878B1 (en) * 1999-09-24 2001-06-26 General Electric Company Method and assembly for connecting air ducts in gas turbine engines
US20040148943A1 (en) 2003-02-05 2004-08-05 Mitsubishi Heavy Industries Ltd. Gas turbine and bleeding method thereof
US6986638B2 (en) * 2002-03-23 2006-01-17 Rolls-Royce Plc Vane for a rotor arrangement for a gas turbine engine
EP1640587A1 (de) 2004-09-22 2006-03-29 Siemens Aktiengesellschaft Kühlsystem für eine Gasturbine, Verdichterleitschaufel und Verfahren zum Kühlen einer Gasturbine
US7246989B2 (en) * 2004-12-10 2007-07-24 Pratt & Whitney Canada Corp. Shroud leading edge cooling
US20090324388A1 (en) 2008-06-30 2009-12-31 Mitsubishi Heavy Industries, Ltd. Gas turbine and cooling air supply structure thereof
US8240975B1 (en) * 2007-11-29 2012-08-14 Florida Turbine Technologies, Inc. Multiple staged compressor with last stage airfoil cooling

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6053701A (en) * 1997-01-23 2000-04-25 Mitsubishi Heavy Industries, Ltd. Gas turbine rotor for steam cooling

Patent Citations (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE3606597C1 (de) 1986-02-28 1987-02-19 Mtu Muenchen Gmbh Schaufel- und Dichtspaltoptimierungseinrichtung fuer Verdichter von Gasturbinentriebwerken
EP0735255A1 (en) 1995-03-31 1996-10-02 General Electric Company Compressor rotor cooling system for a gas turbine
US6099244A (en) * 1997-03-11 2000-08-08 Mitsubishi Heavy Industries, Ltd. Cooled stationary blade for a gas turbine
JPH11125199A (ja) 1997-10-22 1999-05-11 Mitsubishi Heavy Ind Ltd 空気圧縮機ディスクの冷却によるクリアランス制御方法
JPH11315800A (ja) 1998-04-30 1999-11-16 Toshiba Corp 空気圧縮機
EP1074694A2 (en) 1999-08-04 2001-02-07 General Electric Company Apparatus and methods for cooling rotary components in a turbine
US6250878B1 (en) * 1999-09-24 2001-06-26 General Electric Company Method and assembly for connecting air ducts in gas turbine engines
CN1295184A (zh) 1999-11-05 2001-05-16 株式会社日立制作所 燃气涡轮、燃气涡轮机及用于其动叶片的冷却剂收集法
US6986638B2 (en) * 2002-03-23 2006-01-17 Rolls-Royce Plc Vane for a rotor arrangement for a gas turbine engine
US20040148943A1 (en) 2003-02-05 2004-08-05 Mitsubishi Heavy Industries Ltd. Gas turbine and bleeding method thereof
EP1640587A1 (de) 2004-09-22 2006-03-29 Siemens Aktiengesellschaft Kühlsystem für eine Gasturbine, Verdichterleitschaufel und Verfahren zum Kühlen einer Gasturbine
US7246989B2 (en) * 2004-12-10 2007-07-24 Pratt & Whitney Canada Corp. Shroud leading edge cooling
US8240975B1 (en) * 2007-11-29 2012-08-14 Florida Turbine Technologies, Inc. Multiple staged compressor with last stage airfoil cooling
US20090324388A1 (en) 2008-06-30 2009-12-31 Mitsubishi Heavy Industries, Ltd. Gas turbine and cooling air supply structure thereof

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11746675B2 (en) 2021-11-23 2023-09-05 Rolls-Royce Corporation Vane ring assembly for a gas turbine engine with dedicated through-flow vanes

Also Published As

Publication number Publication date
CN102454480A (zh) 2012-05-16
EP2450531A1 (de) 2012-05-09
EP2450531B1 (de) 2013-05-15
CN102454480B (zh) 2015-02-25
JP5411233B2 (ja) 2014-02-12
JP2012097748A (ja) 2012-05-24
US20120114459A1 (en) 2012-05-10

Similar Documents

Publication Publication Date Title
US9416661B2 (en) Axial compressor and associated operating method
EP2060741B1 (en) Turbine arrangement
JP6161897B2 (ja) タービンノズルコンパートメント式冷却システム
US3703808A (en) Turbine blade tip cooling air expander
US7048496B2 (en) Turbine cooling, purge, and sealing system
JP5008735B2 (ja) 蒸気タービン
US9845687B2 (en) Gas turbine engine component having platform cooling channel
US6786694B2 (en) Gas turbine and method of operating a gas turbine
US9605551B2 (en) Axial seal in a casing structure for a fluid flow machine
US10458291B2 (en) Cover plate for a component of a gas turbine engine
JPS61197702A (ja) ガスタービンエンジン
GB2536628A (en) HPT Integrated interstage seal and cooling air passageways
JP2012117540A (ja) 軸流式のガスタービン
JP2007146787A (ja) ガスタービン
US6702547B2 (en) Gas turbine
JP2013249843A (ja) ノズルダイアフラムインデューサ
US5738488A (en) Gland for transferring cooling medium to the rotor of a gas turbine
JP4867203B2 (ja) ガスタービン
US10683760B2 (en) Gas turbine engine component platform cooling
KR101965505B1 (ko) 터빈 블레이드 링 세그멘트 및 이를 포함하는 터빈 및 가스터빈
US9657592B2 (en) Cooling device for a jet engine
US9267513B2 (en) Method for controlling temperature of a turbine engine compressor and compressor of a turbine engine
US10329913B2 (en) Turbine disc assembly
JP7414580B2 (ja) タービン
JP2004036514A (ja) ガスタービン

Legal Events

Date Code Title Description
AS Assignment

Owner name: SIEMENS AKTIENGESELLSCHAFT, GERMANY

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:BENKLER, FRANCOIS;DUNGS, SASCHA;HOELL, HARALD;SIGNING DATES FROM 20111019 TO 20111020;REEL/FRAME:027570/0436

STCF Information on status: patent grant

Free format text: PATENTED CASE

MAFP Maintenance fee payment

Free format text: PAYMENT OF MAINTENANCE FEE, 4TH YEAR, LARGE ENTITY (ORIGINAL EVENT CODE: M1551); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY

Year of fee payment: 4

AS Assignment

Owner name: SIEMENS ENERGY GLOBAL GMBH & CO. KG, GERMANY

Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNOR:SIEMENS AKTIENGESELLSCHAFT;REEL/FRAME:055950/0027

Effective date: 20210228

FEPP Fee payment procedure

Free format text: MAINTENANCE FEE REMINDER MAILED (ORIGINAL EVENT CODE: REM.); ENTITY STATUS OF PATENT OWNER: LARGE ENTITY